text:
Why do airfoil's trailing edges point downward?
The trailing edge of and airfoil is mostly pointing downward. Is there any reason for that? Down wash increases drag and down wash is due to the direction of air pointing down. If the trailing edge is made horizontal (or maybe pointing upward), they can produce less down wash and Kutta Condition can still be satisfied.
What are the specific reasons for this shape?

Downwash is the result of lift (or vice versa, depending on how you look at it).

There is no lift without downwash, so it's usually considered to be pretty important. This gets back to that whole Newton's Third Law thing. A perfectly horizontal wing (at 0 AoA) creates no lift and a wing that deflects air upwards would create negative lift.

@reirab Just to clarify, a perfectly horizontal symmetrical airfoil generates no lift. There are airfoils designed to lift at 0 and even negative angles of attack.

@StallSpin Yes, correct. By "perfectly horizontal wing," I meant that the wing is not deflecting the air stream upward or downward (i.e. it's a symmetrical wing at 0 AoA.)

You are right, the trailing edge does not need to point down. Take symmetric airfoils - here the trailing edge runs parallel to the airfoil chord. Or take reflex airfoils (like the HQ 34 of the SB-13 tailless glider): Here the trailing edge is indeed pointing upwards, and still this aircraft flies.
But to create lift with as little drag as possible, it helps that the trailing edge points slightly down. Why? Because then it will be least in the way of the desired local flow. Lift is created by accelerating air downwards. The wing deflects the air flowing over it, and the trailing edge should reflect this deflection angle.
But - as always - too much is not good: The Eppler 417 airfoil of the SB-7 glider was an extreme layout with too much rear camber. It is said that pilots could still wipe off raindrops near the trailing edge after landing when they had crossed a shower an hour earlier. This means that airflow separated well ahead of the trailing edge on the upper surface, and the raindrops would not be blown away in the separated flow. Flow separation increases drag, and this effect shows that the shape of the Eppler 417 had too much downward camber at the trailing edge.
If the plane is large and heavy, its airfoil's rear camber can be higher - an extreme case is shown below. This is an early supercritical airfoil designed by McDonnell-Douglas, and the highly cambered rear part allows it to integrate very effective Fowler flaps. Flaps help because they allow to change the direction in which the trailing edge is pointing: Low lift coefficients require no flap deflection, or even negative flap angles in gliders, and the higher the lift requirement grows, the more the flaps will be extended, pointing more and more downwards.

The same is true for control surfaces: Depending on the desired lift change, their trailing edge will point up- or downwards. See below for an example where I plotted the pressure distribution for three flap deflections in one graph. Upper and lower surface pressure are shown by color-coded lines, and lower lines belong to the lower surface. Dashed lines show the inviscid pressure, and solid lines the pressure distribution with friction effects added. The wider two lines of the same color are apart, the more lift is created. Note the contour plot below, which follows the color scheme of the pressure plots.

If there are slots to re-energize the flow, extreme trailing edge angles are possible and help to create lots of lift at low speed, which helps airliners to get into small airfields. See below the triple-slotted flaps of the Boeing 727, which was designed for regional traffic from and to small airports.

As usual Peter. I find there's almost nothing left to add after one of your answers :) +1

If lift is produced by accelerating air downward, then higher the downwash higher the lift.
Lift induced vortex also produce more downwash and hence should provide more lift, right?
But, most of the books say that lift induced vortex reduces the effective AOA as downwash is more and hence, lift reduced. How is eff AOA is reduced if there is a downwash?

@Selva: The downwash goes together with an upwash in front of the wing. The suction field above the wing accelerates and bends the airflow - look here for a longer explanation. Maybe this would be a good question of its own.

I will make it as a new question, as it will help others

I just wonder, how low are the flaps upon landing? (Given they're even behind the main gear, I suppose it's extremely low.)

@yo': Do you see the bumper on the rear fuselage? That will make grand contact first, protecting the flaps.

Oh clever! Thanks!

This question discusses different ways to explain why wings generate lift.
The Newton explanation (wings go up by pushing the wind down) can explain why a wing trailing edge points down. This will increase the angle at which the air is directed downwards, and thus increases lift.

I read that ground effect is due to decrease in downwash angle. If as you said, downwash increases lift, then ground effect should decrease lift. But that is not the case. Correct me if i am wrong.

The ground effect combined with Newton's approach of explaining lift is covered here http://aviation.stackexchange.com/a/11396/4197

So, if I understand correctly, ground effect reduces induced drag by reducing the downwash angle.
But at the same time, lift should also reduce, but it is compensated by the cushioning effect (high pressure build up between lower surface and ground)
Am I right? Please correct me if I am wrong

You can see that the resulting force ($F_{res}$) is under an angle (≈ the downwash angle, $\alpha_{eff}$). The vertical component of this resuling force is lift $(L=F_{res} \sin{\alpha_{eff}})$, the horizontal component is the induced drag $(D=F_{res} \cos{\alpha_{eff}})$. The ground effect reduces the downwash angle (therefore putting the resulting force more upright). This increases lift, and decreases drag.

@Selva: You are exactly right.

Down wash increases drag and down wash is due to the direction of air pointing down. If the trailing edge is made horizontal (or maybe pointing upward), they can produce less down wash and Kutta Condition can still be satisfied.

There's a very important reason why we don't want to try to decrease the amount of downwash the wings generate: the amount of downwash is equal to the amount of lift!
An airplane in straight and level flight must produce an amount of downwash equal to its own weight. If it produces less downwash than that, it will fall. There's no way around it.

Which way the trailing edge "points" is related to the in-flight angle-of-attack, not the way someone might have arbitrarily drawn a section picture.
If you've ever seen an aircraft flying upside down (relative to airflow), even then - the trailing edge is still pointing down, despite the inverted wing.
It always points down, because that's the only way to generate the lift needed to counter the weight of the aircraft.  Even if you invented your own trailing edge that pointed sideways or up, the airflow itself will still need to be "going down" after that obstruction you put there, if the craft is going to fly.
The shapes of wings is usually optimised using Navier-Stokes to find shapes which produce lift for little drag - and those equations produce the best results with the trailing edge downwards.


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How is heavy cargo secured in military cargo carriers?
Some big military cargo carriers carry tanks and other heavy equipment and I am not sure if they have some kind of special locks or tie-downs in the cargo space to manage those heavy loads.
How are these very heavy loads secured?

You mean tiedowns?

i guess tiedowns are for the complete aircrafts but do we use the same inside the cargo as well as to tie the tanks etc?

@NitinG - I've tweaked the question a bit, I hope this retains the flavor of what you were asking, while clarifying it.

Not always well enough, unfortunately. http://www.ntsb.gov/news/press-releases/Pages/20150714b.aspx

^ Accompanying video of the crash: https://www.youtube.com/watch?v=-MB9JDBe4wA

The cargo in a military aircraft are secured in the cargo hold using chains, straps etc. connected to stowage tie-down devices, which go into tie-down rings . The following image shows a M1 Abrams tank secured in C-17 Globemaster III using tie-down devices.

USAF uses three different types of devices to secure cargo in aircraft, which vary in their load capacity:

MB-1 tie-down device with 10,000 lb rated capacity.
MB-2 tie-down device with 25,000 lb rated capacity.
CGU-1/B tie-down device with 5,000-pound rated capacity

The MB-1 and MB-2 are similar in their form and function, with the only difference being their load capacity.This image shows a close-up of the MB-1 and MB-2 tie-down devices:

"C-17 Globemaster III no. 5139 stowage tiedown devices" by BrokenSphere - Own work. Licensed under CC BY-SA 3.0 via Wikimedia Commons.
The hook goes into the tie-down ring, while on the other end, straps or chains (from the cargo) go.
The CGU-1/B is a 20-foot nylon web strap with two metal hooks at both ends. While one hook is stationary, the other hook has a ratchet device and can be moved and tightens the device when it is being used.

Note: Apparently, Military designation system refers to MB-1 and MB-2 as CGU-3/E and CGU-4/E respectively.

In the Abrams picture it looks like one of the chains has a lot of slack. Do you know whats up with that?

I think what you're looking at is the slack unused length of the chain used in the tie-down. It's wrapped over the taut span above it which is the same chain.

In other words, they use the same things to tie the equipment down in the plane as they would on a ground based transport. They probably just throw an extra one or two in because a shifting load on a truck is an annoyance while in a plane, a shifting load can be fatal.

@FreeMan: Load shifts in a truck can be much more than an annoyance, and can even be fatal as well. Here's one: http://wyandottedaily.com/fatal-accident-may-have-been-caused-by-load-shift/

Load shifts are problematic no matter the mode -- a shifted heavy load on a flatcar can derail a train or worse yet, wipe out a through-bridge structure altogether; shifted loads can also cause stability hazards in marine transportation.

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text:
Which is longest Oceanic crossing for any commercial flight route?
During oceanic crossings, there are limited to no terrestrial radio communications available, so I would think that most flight routes are designed to minimize ocean crossing time.
Which commercial flight route have the longest oceanic crossing?

Most of the flight routes are designed to minimize fuel burn and if it takes the plane over the ocean, so be it. Communication is restricted, but HF radio works and so do satellite links.

Related: Are there diversion points for southern Pacific great circle route flights?

As a sidenote there is notion called ETOPS which may modify the optimal flight path (along with the weather), depending on the aircraft and airline and they may need to use different routing because of ETOPS certificate

Note that on transoceanic routes, airplanes use HF radio, which bounces off the ionosphere and thus allows it to reach over the horizon for very long distances. Most modern airliners also have satellite links. So, they are free to fly whatever route works best for weather (within ETOPS limits).

For the longest  Oceanic Crossing, the following are the top three contenders:

Sydney to Vancouver
Melbourne to Los Angeles
Auckland to Vancouver

Out of this, 1 and 2 pass over a few islands on their way, 3 does not pass over any landmass.
Have a look at this : www.gcmap.com
Image :

(Image Source: www.gcmap.com)

It's worth noting that the actual routes flown by aircraft reflect current weather patterns, and can vary considerably from the great circle path. For example, AKL-YVR frequently goes west of Hawaii.

The Air Canada Melbourne-Vancouver is probably the longest. Quibbling as to whether it might pass over an island is pointless!

Currently not operated but there are longer routes such as CPT-JFK (Cape Town to New York) 7,797 mi OR SCL-NRT (Santiago to Tokyo) 10,666 m

I do wonder if JNB-ATL (currently flown by Delta) would be in the running - 8400 miles but some of that is over land.    Within South Africa, Cape Town is better situated for this, and Delta and United are soon starting to fly ATL-CPT and IAD-CPT respectively: https://www.routesonline.com/news/29/breaking-news/298357/delta-and-united-set-to-secure-cape-town-rights/

ATL-CPT is 8130 miles (Atlanta is perhaps 250 miles inland) and IAD-CPT is 7925 (IAD is about 150 miles inland).  By comparison YVR-SYD is 7757.

The Qantas Sydney-Santiago service comes in at 7,060 miles (11,362 km) and certainly crosses some very remote stretches of ocean - http://www.gcmap.com/mapui?P=syd-scl%0D%0A&MS=wls&MP=p&DU=mi

There are also SIN-LAX or SIN-SFO handled by Singapore airlines or United. The flight time is around 17h. Singapore to New York would be even longer, but I am not sure if they fly over Pacific Ocean...

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How does the stick pusher prevent an aircraft from stalling?
The stick shakers warns the pilot of stall conditions and some fixed winged aircraft have a stick pusher which tries to avoid the stall. What is the working methodology of stick pushers?

A simple explanation is that a stick pusher is a:

... device that pushes forward on the elevator control system
  whenever the aircraft’s angle of attack reaches the pre-determined
  value, and then ceases to push when the angle of attack falls
  sufficiently.

A detailed explanation is:

A stick pusher system incorporates both a passive monitoring component
  and an active intervention component that consists of an hydraulic or
  electro-mechanical pusher mechanism. The monitoring component
  constantly evaluates critical flight parameters such as airspeed, load
  factor and flap setting against the aircraft angle of attack. If the
  angle of attack reaches a predetermined value, the intervention
  component of the system is activated and pushes forward on the
  elevator control system until the angle of attack is reduced to an
  acceptable value. The system operates autonomously and without any
  required action by the pilots.

FAA has an advisory circular (AC) for stick pusher training.
For Embraer 145, this video shows a demo of stick shaker, stick pusher and the warning systems.

To elaborate a bit on what the other answers say we must discuss what a stall is and how it happens.
First off a stall can happen at any speed or bank angle
A stall by definition:

...is a reduction in the lift coefficient generated by a foil as angle of attack increases (see picture for what happens to the air when you stall)

and by definition a stall

...occurs when the critical angle of attack of the foil is exceeded

So in an aircraft the typical way to recover from a stall is to pitch the aircraft down. This helps to gain speed but mainly reduces the angle of attack of the wing (the angle you exceeded the limit of to stall in the first place) which in turn helps the wing regain proper air flow over it and begin to generate lift again. Thus a stick pusher servers to automate this function of pushing the stick forward (pitching the plane down) to recover from the stall. This can also help in situations where its difficult to know if the plane has stalled much like the accident that occurred on Air France 447

In the case of AF447 it is doubtful Bonin would have reacted any differently to a stick pusher than he did to the 75 loud "STALL" warnings which he ignored. If he didn't trust the audible "STALL" alerts, he probably wouldn't have trusted any other manifestations of the stall detection systems. Reading about Bonin's actions it seems to me unlikely a stick-pusher would have made much difference.

nb Colgan AIr Flight 3407 "As the aircraft came even closer to stalling the stick pusher activated ("The Q400 stick pusher applies an airplane-nose-down control column input to decrease the wing angle-of-attack [AOA] after an aerodynamic stall"). The captain overrode the pusher and continued pulling on the control yoke causing the aircraft upset and later loss of control."

Regarding methodology here's how Douglas does it on the MD80.

The stick pusher operates only with slats at mid or extend. It comes on 6 seconds after lights and horn or when AOA increases 4°. If auto slat has failed, stall lights and horn are 0.2 seconds after stick shaker, and the 6 second to override request only 2.5° of AOA for stick pusher.


text:
How long does it take to train a commercial pilot to fly a fighter jet?
The question is motivated by the recent events in Ukraine and the Ukrainian parliament on Monday tweeted that Europe was sending 70 fighter planes to Ukraine.
If this were to happen, how much time would it take, roughly, to train a commercial pilot to fly a fighter jet? (provided that the pilot is in excellent physical condition)

The linked and somewhat nebulous article suggests the 70 jets are all models similar or identical to those in Ukrainian service, flown by existing pilots. Suggest edit to define if this question is about this type transition process, which is pretty common and probably has good sources, or the more nebulous 'how long to train an F35 pilot to some arbitrary level'.

Not an useful answer, but with enough twisting you could claim 12 minutes to learn to fly a fighter https://en.wikipedia.org/wiki/Holden%27s_Lightning_flight

@GremlinWranger The "is this were to happen" meaning that, in the future, Ukraine might need new jet pilots as they could be "running out" of the current ones (could not find a better way to say that, sadly). And since they can't have foreigners. they might have to train pilots from other types of planes, like the commercial pilots.

There's a fairly high chance that a number of commercial pilots were fast jet pilots before becoming commercial pilots as thats the more usual path to commercial aviation.

@Jamiec In the US, but in a smaller country like Ukraine? That would be interesting to know...

@deju vu - the difficulty in answering the question for Ukraine is that it becomes a political one. How acceptable is it for pilots to die, loose aircraft or hit friendly forces due to lack of training?

This doesn't answer the question, but, with regards to what prompted it, the types of aircraft that were said to be in the process of being given to Ukraine are ones that Ukrainian pilots have been flying since the 1970s. They're very accustomed to these aircraft. Ukraine does have an Air Force. It's not like they would be sending airline pilots to fly them. That being said, it appears that there was some significant miscommunication. The countries that were said to be donating the aircraft have now stated that no such agreement existed, as per updates to the article linked in the question

Just my guesstimate: Even if such fighters were being donated (how many, in that case, is unclear), I haven't seen information whether they still have radio and other equipment compatible with what the Ukrainian Air Force is using, or whether they have been modernised to NATO standards. If the latter, they would need to be un-modernised to be useful.

@Déjàvu in regard to the usual carrer path - in post-Soviet countries becoming a civilian pilot without becoming a military pilot first was almost unheard of. On the other hand, not all military pilots are a fighter jet pilots.

In almost all modern air forces, number of trained pilots typically exceeds number of available aircraft at any point in time.  Historically there have been times where this hasnt held true, such as the Battle of Britain where production exceeded training output, but complexity of aircraft has gone up to the point where maintenance is an issue these days.

The donated aircraft would be flown in by pilots who can fly them, so maybe you could ask those pilots to stick around for a while and fly them into battle... why the hurry to leave?

@GremlinWranger Holden took three attempts to land, had a tail strike, broke the parachute, and burned the brakes before stopping 100m from the end of the runway.  I don't think you can call that "having learned to fly"  He survived, but it wasn't a success.

fly it or fight it? There is a major difference.
A few days studying the systems and maybe a few hours in a simulator and/or a two seater with an instructor and he could probably fly the aircraft and take off and land without damaging it seriously enough to need major repairs (depending on the type of course, some are notoriously difficult to handle).
But to operate it effectively as a weapons platform so he's not just flying around in circles while people take potshots at him takes a lot longer. Months, at least in peacetime, if not years.
That's the drawback of how advanced and complex machines have become since 1940 when raw recruits with a few hours of instruction were thrown into the Battle of Britain and expected to defend the realm against the Hun, and succeeded (though at very high cost of life and aircraft).

I still remember that day some 30+ years ago when as an engineer in the USAF, I was introduced to the reality of a fighter pilot. I was watching a cockpit video from an F-16 while a pilot was narrating how the radar intercept was progressing. The pace was similar to an announcer at the horse races. At some point I asked; When is he flying the plane? The response was; Oh, that's a background task. That's when I realized that an F-16 isn't a plane, it's a weapon system that just happens to fly.

Everything changes when "Aviate, Navigate, Communicate" adds in "watch for bad guys, identify your mission target, don't get shot, select the correct munitions, configure and target them correctly, evade counterfire, coordinate with air and ground forces, etc, etc, etc...".  Just keeping situational awareness is a monumental task by itself.  Consider, for example, how much a bus driver would need to learn before they could effectively command a tank.

The pilots of SR-71s were 'a little busy'. Seated behind them, the reconnaissance systems officer did everything else.

@Mazura And the SR-71 isn't even a warplane - no dogfights, no weapons; just cameras, really.  Some multirole/fighter class planes did have a RIO/WSO, of course.  In an airliner you always have at least two flight crew also and they don't have any extra stuff to worry about at all.  Instead of two pilots managing flight you have one who has to play PF and PM at the same time.  For one-seaters without the luxury of a RIO/WSO you're now down to just one pilot who is PF, PM, and RIO/WSO all at the same time.  The workload is crazy.

@J...: And yet, pilots of single-seater fighter jets like the F-16 joke about rather having the extra gas than another person in the cockpit with them.  e.g. https://youtu.be/REtqz-zIhP4?t=126 (Mover Ruins Movies: IRON EAGLE (1986)), he says "that's the way it should be" about being all alone in a single-seat cockpit.  He flew F-16s for the US air force reserve, and F-18s for the navy, and now T-38 adversary air for the air force.  If I spent longer looking, I could probably find something less jokey, maybe in one of his interviews on his channel with an F-14 pilot about their RIO.

(My impression is that older jets with less nice radar / weapons electronics took more attention to get the most out of them, so it made more sense to have a dedicated person managing it.  And also less advanced fly-by-wire meaning that managing the plane itself took more attention?  So single-seat would have been even more of a problem for jets with long-range missiles?)

This short training looks sufficient for a kamikaze flight.

@PeterCordes increased automation has reduced the need for a GIB (guy in back), but it's still there for specialised stuff. Hence the EF-18 Growler being a two-seater, and the Strike Eagle, for example.

What's the reference to "the Hun"? There's no mention of "the Hun" on the WP Battle of Britain page.

@AzorAhai-him- that's a colloquial reference to the attacking Germans.

@J... While the SR-71 didn't have weapons that doesn't mean it didn't have to worry about incoming weapons.  Various places that didn't like eyes in the sky took potshots at it now and then and the crew had to spoof or outrun the incoming missiles.

@LorenPechtel Usually outrun.  I believe SOP on the 71 was just to accelerate if fired upon.  It's not like you can do anything aerobatic at FL810 and mach 3.2 without falling out of the sky and missiles don't have the endurance to catch up.

@J... It's pretty hard to outrun a SAM fired from in front of you.  You're going to have to dodge or jam such a missile.  (Note, however, that most SAMs are pretty easy to dodge at that point--the air is so thin their control surfaces aren't very effective.)

As a baseline answer, the Royal Australian Air Force has a rundown on the recruiting page, hidden under 'full details' option and 'training' where after completing initial flight training they have:

Pilot Advanced Courses (1 year). If you are streamed to go down the FJP pathway you will be required to complete the following courses:
Pilot Advanced Course (6 months). Remaining at RAAF Base Pearce you will undergo training on the Hawk Lead-in fighter at Number 79 Squadron (79SQN).
Lead-in-Fighter Course (6 months). Once you have learnt the basics of flying the Hawk and operating at higher airspeeds, you will be posted to 76SQN at RAAF Base Williamtown to learn how to operate the Hawk in a tactical environment (air to air and air to ground).

This is training on the Hawk, a fast jet with weapons capability but not a current combat aircraft. Then they have

Operational Conversions (6 months). At the end of the advanced courses you will be streamed for operational conversion onto one of the following platforms:

And then listing the F18F, F18G and F35
So the starting point for a hypothetical pipeline for pilots to effective fighter pilot is 18 months. While a war footing might let you strip weekends and other down time out it seems unlikely you could do better than 6 months. In practice I would suspect that the key factor in this timelines is training assets - do you have enough simulators, instructors, training aircraft and a safe places to fly for all your pilots, for Ukraine I'm sorta expecting the answer to most of above is 'no'.
So if you need highly effective general purpose pilots you need to have started a year or so ago. If you are at the point of sacrificing one or more of 'effective', 'general purpose' and 'pilot' then other things become possible.

For Ukraine, I'd guess that providing training to potential pilots is something other countries could help with.  i.e. fly the pilot trainees to another country where they can train.  Not that it would make sense, because the timeline for becoming combat-effective is just so long.  Modern jets are too expensive to risk on half-trained pilots, and full training takes a long time.

At the end of January, 2023, Ukraine said that about six months are required to train they pilots for combat in western fighter jets such as F-16s.
Ukraine previously said they have some pilots who have already completed this training.

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text:
Is the size of hail on the ground any indication of the winds and updrafts of a thunderstorm?
Is there any rule of thumb or way to estimate the wind speeds or up and down drafts inside a storm by the size of hail that reaches the ground?

The October 2018 issue of AOPA's "Flight Training" magazine provided the following illustration and comments.

A 24mph updraft will form ~1/4" hail, while a 98mph updraft will form
4" hail.
Down drafts are usually about 1/2 the speed of the updraft
(due to a larger outer core area).
When a pilot encounters a
downdraft, expect an updraft twice the speed to soon follow.
The change will be a quick dramatic shear.
Only about 10% of thunderstorms produce hail.

Certainly. There are two effects involved here which I will describe briefly as follows:
first of all, a small spherical object will exhibit a well-defined terminal velocity when free-falling in air. This means for a hailstone to have formed at a given diameter, the updraft velocity would necessarily had to have been close to the terminal velocity for an object of that size, for it to have stayed in the storm cloud and grown to that diameter.
Second, if you section a large hailstone you will find that it is layered, like an onion. this means that hailstone contains evidence of not just one freezing event but instead as many as there are layers in its structure.
In each case, then, the velocity of the updraft it experienced at each individual stage of its growth had to be significantly greater than its terminal velocity at that stage in order for it to have been lofted back into the storm cloud where it then got another ice coating before it began falling again.

Current research debunks the (somewhat commonly held) notion the hail cycles through the updraft multiple times, but rather makes one trip one and then falls to the ground and does all its growing and accretion within that single trip.


text:
At what Reynolds Number does laminair airfoils stop working as expected?
I was watching a RC Plane builder on YouTube create an airliner on a 3D printer and noted he used the scale laminar airfoil of a B-727. It seems to me that laminar airfoils don't work at the scale of bugs, probably don't work for birds, and possibly don't work for RC models with less than 2-3ft wing spans.

Intuitively, it seems laminar airfoils are tied to Reynold numbers. Is that assumption true?
At what Reynolds number or scale (or other limiting factor) are laminar airfoils no-longer operating as expected?

The B-727 does not use a laminar flow airfoil.  It may use a supercritical airfoil -- which may get some limited amount of laminar flow, but nobody would consider it a laminar flow wing.
Any body or airfoil designed for laminar flow will have a somewhat narrow operating range where it has favorable characteristics.
Here is a modified figure from this paper on the design of bodies of revolution for natural laminar flow.

Each of the five bodies was optimized for its particular region of Reynolds number.  Notice that the shape of the bodies is quite different.
For each one, going above its designed operating range is very bad.  Operating below its designed operating range is not catastrophic, but better performance would be possible with a body designed for that (lower) Reynolds number.
Note that the drag coefficient and Reynolds number have been defined in terms of the volume -- to try to keep the comparison of these bodies as 'apples-to-apples' as possible.

Yes, certainly Reynolds # is a key parameter effecting airfoils designed to have laminar flow.
As a rough guide, for Reynolds #’s < 100,000 the boundary layer is most likely laminar. By Re ~ 1,000,000 the boundary layer is usually fully turbulent. Some factors that can effect the extent of laminar flow include noise, pressure gradients (including shocks), crossflow instabilities (on swept wings), surface roughness or waviness, and wall boundary conditions (e.g., thermal, suction, blowing).


text:
Why are general aviation avionics so expensive?
I am quite sure that someone already asked this question, but I couldn't find it on this site.
Why are avionics for general aviation aircraft so expensive?
For example:

The Garmin® G5
$ 2,149.00

The mighty Garmin® G1000
$ ~ 50,000.00

Nobody can tell me, the hardware and development for a small screen costs more than 2,000 dollars. Or two screens with some knobs for the price of a middle-class car. (Of course, those avionics are not bad, they are worth every cent. But it still looks expensive for me.)

What causes this, quite high price?

I guess it's because of high licensing costs to license it for use in aircraft, or are the production costs really so high?

IMHO - the answers are "scaling" and "certification".

Related questions: here, here

"Nobody can tell me, the hardware and development for a small screen costs more than 2.000 dollars.".  Why not?  They can, and they do.  This is not consumer electronics produced in hundreds of thousands or millions of units.

Compare a non-TSO instrument with a TSO instrument and you'll understand where a good portion of the costs come from. These two units are exactly the same, almost a $1000 price difference.

Hardware? No. Development? Definitely. The effort spent on developing the hardware and software is counted in man-millennia. Times at least $100,000 per man-year and divided by the mere tens of thousands of units produced…

(Full Disclosure:  I work at a company that makes avionics, although I work in a different, unrelated division.)  Its amazing how much work the avionics software developers have to do to make sure every line of code is rock solid.   Every tiny piece of code is absolutely perfect, has been tested an absolutely insane amount, and performs incredibly fast.  This is nothing like a "regular" Android App, which might all manner of bugs or slightly "odd" behavior.  They make every effort to make the software fast, rock-solid, and prove that it is solid.

@abelenky - Quite true. If your Android app crashes your phone, no one gets hurt. Avionics software is specified, designed, and verified based on the criticality on its failure. Of those 3, the design/implementation is maybe 20% of the effort.

If you want to consider a similar regulatory industry look at implanted medical devices. There is also a good reason that the company making 3D printed bionic arms is in the UK and not the US.

It's interesting to compare the dedicated and refined look and feel of the images in your question to something like DIY avionics for experimental aircraft

@cwd This is really cool. Reminds me of my car full of smartphones and a Head-Up-Display to mimic all the cool stuff modern cars have.

There are a few reasons,

As many have mentioned the FAA's rigorous certification process is time consuming and complicated to navigate. You can find a full overview from them here. In short you need to do a lot of paperwork and thoroughly test and prove your designs.
Once you have your TSO/STC/TC on the product you must maintain it. This means you need to budget in for future problems and updates. If a plane goes down as a result of your avionics and the FAA decides to ground all aircraft that have your part installed (rare but they can do it) you are going to have a lot of angry customers on your hands. While I do not believe a company is legally obligated to provide a fix it is in their best interest if they wish to stay afloat.
Charts change all the time, most modern avionics with navigation capability require regular updates. This is something the company needs to maintain for their avionics to be legal for certain kinds of operations.
The hardware is expensive! avionics generally require a greater level of precision than your iPhone GPS and 3 axis sensors. Most modern AHRS use micro-mechanical internals, magnetometers and other complex (read costly) units. These units need to maintain precision at all times. Generally speaking they are also double or triple redundant so you are paying for 3 of them while an iPhone may only have one GPS receiver. The screens are also high quality, daylight readable screens, small additions but it all adds up.
Like in any high speed/high risk industry I am sure the insurance premiums to insure an avionics business is high. Keep in mind that the FAA/DOT has the ability to blame you the manufacturer in the case of an incident that was the result of a faulty part.
Avionics (glass in particular) can be certified for specific air frames. This means that you may need to work with multiple companies to test out your platform in multiple air frames and program in multiple engine/air frame data sets. You may also need to build bracketing to mount and install the system. Some systems are designed to be drop in replacements like the Aspen units. Others may require larger panel overhauls to install. There are a huge number of other systems aside from the air frame/engine, like the auto pilot, other radios, ADF units etc. that your system may need to interface with. You would need to test with all of them to include them on the AML list of things you can install the unit with. This is a costly testing proposition in and of its self.

The other thing to consider is that even with your TSO, a Type Certificate or Supplemental Type Certificate likely needs to be obtained.

Edited to reflect

Rule of thumb: about two-thirds of the price of anything you buy for a GA aircraft goes to liability insurance.

You forgot scarcity... in the sense that only a few companies are certified and/or make quality avionics people want to buy (so they can charge whatever they want -- captive market sort of thing), and also in the sense that there's not exactly huge amounts of these things flying off the shelves every day (low volume of sales), which drives costs up per unit for the manufacturer, which means it drives costs up for the consumer.

"Hardware is expensive".. really, no.

Other answers cite certification, and that is a big cost.  Additionally, a very large issues is the number of units that will be sold, and supported for the unit lifetime.
For comparison, Samsung has shipped  10 million S8 Galaxy phones within months of release.  While the total size of the US General Aviation fleet is  under 250,000.  How many of those aircraft can Garmin realistically hope to put the G5 in?  Avionics are rarely upgraded, and aircraft are kept operational for a long time.  If the development costs are at all similar, but Garmin expects to sell 4 orders of magnitude fewer units, they need to charge more of the development cost for each unit.

The cost of rigorous FAA certification versus number of units produced would be my subjective opinion.  Count on our little glass horizon not full up Galaxy 7 catching on fire or even just failing when needed.  That's one of several things we pay for, again: opinion without data.

Basically, the development and certification costs are high and the production volume is low. So each unit that is sold has to pay for a substantial part of the development costs.
For comparison, Apple's iPhone's are produced at a production rate of over 500 000 per single day. There is no way any avionics product will ever be produced in such a volume over the entire product lifecycle. For many avionic products a total volume of 10 thousand is a number that will never be reached. Compared to an iPhone that is procuded in a volume of approximately 200 million units, the factor of cost for development in the price is 20 thousand times as big .

I now found what I think is the answer:
Wikipedia: Avionics software
It's really a great expense to develop an avionic.

The main difference between avionic software and conventional embedded software is that the development process is required by law and is optimized for safety. It is claimed that the process described below is only slightly slower and more costly (perhaps 15 percent) than the normal ad hoc processes used for commercial software.

The process of development is really huge and the price is justified.

Human interfaces
Hazard analysis
Maintenance manual
Design and specification documents
Code production and review
Unit testing
Integration testing
Black box and acceptance testing
Certification

I can't imagine how this must have been looked while developing the G1000 or similar.
Source of quotes

I expect you will find that the added expense for using DO-178B/C rather than general industry standard processes increases cost much more than 15%, especially for Level A or B software. I've seen estimates of upwards of 50% for Level A software.

I believe this is basically the correct answer.  You can't hire a competent programmer for under 200k, and if you want "the best" it's pricey.  Shit software is quick to develop, solid embedded-like software is tough and slow to develop.

You can certainly hire excellent programmers for under 200K (and even 100K). And embedded software isn't intrinsically more difficult or time consuming. The added cost is really in the certification burden placed on avionics.

"The process of development is really huge". What you posted is valid for every software development (except for certification). As a (competent) programmer, I can tell that coding avionics presents no special difficulties.


text:
Is this a wing-tip device on the A-10 "Warthog"?
Source: Flickr courtesy of: U.S. Air Forces Central Public Affairs, photo by Master Sgt. William Greer
Notice at the side edges, the wing peels down and back around towards the fuselage. (The airplane is upside-down in this photo.)
Never noticed this before. Is this a wing-tip device producing more lift? If so, what kind is it? I would like to read more about it.

Thank you for the gorgeous photo. The detail behind the Warthog, especially the snow line, is absolutely phenomenal.

It is a great photo.  Could you add attribution?

Looks like this might be the source: https://www.flickr.com/photos/dvids/5510306058 *> An A-10 Thunderbolt II from the 23rd Fighter Group, Moody Air Force Base, Ga., peels away after being refueled from a KC-135 Stratotanker, assigned to the 340th Expeditionary Air Refueling Squadron, while flying over Afghanistan in support of Operation Enduring Freedom, Feb. 26, 2011.
U.S. Air Forces Central Public Affairs
Photo by Master Sgt. William Greer
Date Taken:02.26.2011
Location:SOUTHWEST ASIA, AF*

Reference a Civilian (737) question here:  http://aviation.stackexchange.com/questions/21122/what-are-some-differences-that-make-the-new-737-max-winglets-more-efficient-than/21127#21127
Civ. aircraft look much harder at fuel economy, etc. but it seems the primary benefit is low-speed performance and better overall performance.

The wingtip devices used in A-10 Warthogs are called drooped wingtips (also called Hoerner wingtips in some cases), which essentially increase the aspect ratio of the wing by forcing the vortices further out.

Source: zenithair.com
There are a few reasons for having this wingtip device:

The drooped wingtips act in a manner similar to the winglets and reduce the induced drag. As a result, the loiter capability of the aicraft, an important one for Close Air Support (CAS) aircraft, is improved. The A-X competition, that lead to the A-10 develpment, required the aircraft to have Two hour mission loiter time at max mission radius with 9,500 lbs payload
The drooped wingtip increases the local span loading near the wingtip, due to which the aileron response is improved, increasing the manuverability at low altitudes. The A-X competition required the aircraft to be Highly maneuverable below 1,000 ft
The drooped wingtips also improve the takeoff performance, an advantage as the aircraft is expected to be operated from short, forward fields. Infact, drooped wingtips are found in STOL kit of some aircraft, like DC-2 Beaver. One of the requiremens of the A-X competition was that the aircraft should have 4,000 ft takeoff distance at MTOW.

As can be seen, the drooped (down) wingtips wingtips serves these purposes.

There seems to be some confusion regarding the reason for these wingtip devices. One theory (@Peter Kämpf) is that they are there to protect the ailerons from ground contact. However, I think this is unlikely, given the ailerons, when extended, go well beyond the wingtips, as shown below:

Source: scienceforums.net
Dropped down wingtips do serve that purpose (sometimes) in gliders. However, there is no reason to believe that is the case here, considering the height above the ground and the aspect ratio. The drooped wingtips, on the other hand serve to hold the countermeasure system.

Source: rcgroups.net

The improved take-off performance isn't really a reason to install drooped tips, other types of winglets would also improven the take-off performance.

@ROIMaison - The drooped wingtip, unlike other winglet styles, acts as somewhat of a pressure barrier, keeping air under the wing in addition to modifying the tip vortex. This improves lift at high AOA such as during takeoff rotation.

This comes at the cost of reduced reduction of vortex wake compared to an upward-swept winglet, which for the A-10 was a secondary concern as range, speed and ceiling requirements didn't require ultra-low-drag (the thing can carry over a dozen Maverick missiles; wingtip vortex drag is not the first item on the list for drag reduction)

A reduced reduction of vortex wake? Does that mean it makes more votrex wake, @KeithS?  :)

@FreeMan I read that as "more vortex wake than an upward-swept winglet, but less vortex wake than no winglet at all"

The picture you're using to show the reach of the ailerons has them split and fully deflected to be used as speed brakes (something Peter Kämpf also mentioned), and they are possibly well beyond the angle of deflection achieved when they are closed and being used as an aileron.  The actual deflection required for the aileron for maneuvering when in a bank (which is the suggested usage case) seems to be covered by that wing tip bend, so Peter's idea still holds up.  I'm not saying you are necessarily wrong...but the evidence you are using doesn't actually support your argument.

@JayCarr If you're in risk of your ailerons contacting ground while in a bank in normal flight, there's plenty of other things hanging off the bottom of an A-10 which will contact first.

I just read on the Wikipedia page for the Cessna 172: "The wings can be modified using a number of STOL modification kits, some improving high speed/cruise performance but most concentrating on STOL performance. Horton's STOL kit is one of the better-known of the latter. It involves fitting a more cambered leading edge cuff to increase the maximum coefficient of lift, fitting fences at the aileron/flap intersection and fitting drooped wingtips."

On the Horton's page, they claim this: Conical Cambered Wing Tips -
Low speed aileron effectiveness is improved by special wing tips. Induced drag is further reduced since the HORTON STOL-CRAFT tips reduce wing tip vortex strength. They also increase the wing area and decrease the stall speed.

If I add raised wingtips to one of my RC planes they significantly decrease  the aileron authority, meanwhile improving the dihedral effect and the self-stabilizing tendency. But for maneuverability, it's not as good. If I add an autopilot, like Ardupilot, the raised wingtips will make harder for the autopilot to instantly correct the leveling when passing through turbulence, because the effect of the ailerons is dampened by the self-leveling tendency. Basically, the ailerons have to fight not only to change the attitude of the wing, but against it's tendency to rotate by itself, that might not act in the right direction when flying through turbulent air.
If I add dropped wingtips, they improve lift (the same wing is capable to carry more weight) and they increase greatly aileron authority, while decreasing passive auto-leveling.

The plane becomes harder to fly manually, but for the autopilot the effect is good, because the smallest deflection of the ailerons will have a big effect. The autopilot is capable to correct instantly the horizontal attitude of the wing, by moving the ailerons very fast, and in minute increments. The plane will look like it's flying on rails, even when it passes through turbulence. And the lower the altitude, the greater the turbulence, because the wind is moving over ground obstacles and it's not flowing laminarly. When the wind is strong and there are bushes and trees on the terrain, the turbulence can be so great, that a RC model can become impossible to fly low if it doesn't have some sort of electronic self-leveling device.
The conclusion is that, for a plane that has to fly low, dropped wingtips and an active self-leveling device are the best combination. If the plane was low wing - high CG, I would have used raised wingtips.

Interesting answer from another perspective, welcome to Aviation.SE :)

Indeed! Excellent answer with practical experience that directly applies to the A10's flight profile.

Can you post a picture of your RC plane? Or at least tell us it's basic shape? Are the wings high-aspect ratio rectangles like the A-10? I'm interested because there are plenty of raised wingtips on modern large commercial jets, which have swept wings around 30 degrees. Never heard of any poor handling characteristics this incurs at low speed low altitude, so it could be the swept wings.

Commercial jets have very different needs. They fly high and have to be very fuel efficient. They have a low wing, which by default has an anhedral effect, and raised wingtips annihilate that and make the plane neutral, that's why they have good aileron authority. The situation of the commercial jets is more complex. their wings are aerodynamically very elaborate, with a lot of slates, flaps, spoilers etc that can change the configuration of the wing for every needed speed and attitude. So it's not fair to compare such complicated monsters with simpler wings.

It's called a drooped wingtip.
The general idea is to move the vortex away from the wing, reducing it's influence.
As with all the winglet variants, opinions vary about whether or not it generates more lift than other variants.

Source
There are more aircraft that employ this type of wingtip device:

Source

Source

That is odd. I thought moving the vortex to a spot above the wing surface was ideal, because that low pressure vortex can add lift that way. Moving it farther away from the wing (spanwise) doesn't seem to change anything that I can see.

See this answer http://aviation.stackexchange.com/questions/21799/why-does-the-vortex-created-by-wing-affects-its-own-angle-of-attack/21803?noredirect=1#comment49685_21803. Moving it further away reduces its strength and effects.

I'm guessing they don't work as well as a winglet.

@DrZ214: Moving it away from the actual wingtips have the effect of reducing the vortex size. Up or down doesn't seem to matter much.

@slebetman, why would up or down be any different from sideways? The velocity is only a function of distance from the center:
https://en.wikipedia.org/wiki/Vortex

@ROIMaison: Yup. Increasing span also works. In fact, it's the best option. However, if you're span limited, winglets or winglet-like devices makes sense. It's just that up or down doesn't matter much.

Oh boy, you guys are obsessed with that wingtip vortex, seemingly the source of all drag. No, the reason for this wingtip shape is much simpler.
It's about the protection of the ailerons from ground contact at low level flight. This is especially important for gliders with their central wheel, but also for aircraft which maneuver a lot at low altitude. They will benefit from something that takes the impact forces and keeps the control surface intact and working, especially when this surface will be deflected downwards at the time of ground contact.
Honestly, this is the reason, not some esoteric vortex shifting.

The right wing of this image http://www.foodman123.com/cut42.jpg, shows that the ailerons stick out quite a lot further downwards than these tips do, so then the ailerons (if deflected by a reasonable angle) would still hit the ground ealier than the tips?For the gliders it makes sense, but for the A10, it's a bit out of scale. If they wanted to protect the ailerons, to me it would have made more sense to place a larger protecting tip.

@ROIMaison: The ailerons of the A-10 split when used as a drag device. What the drawing you linked to shows is the deflection when used as a speed brake, and yes, with that it will touch the ground before the wing tip does. Better read this. Also, the aileron doesn't quite reach out to the tip and with some bank the wingtips give sufficient protection.

Thanks for the link, it is an interesting read! In this document (p43) I found the following phrase:
"The wings have Hoerner wingtips that reduce induced drag and wingtip vortices. They also improve aileron effectiveness at low speeds." so it seems they also agree on the vortex reduction effect.

Also, consider this image. The hardpoints on the wings provide plenty protection for control surfaces, it seems. I'm afraid this answer just doesn't apply to an A-10.

@sanchises: For one, this A-10 is clearly at rest, so aileron deflections will not help. Also, it is not banking. Why would you think this picture applies here?

@ROIMaison: This is not the official manual. It is a fan product (citation: The manual for Eagle Dynamic's DCS: Warthog). But it contains a lot of details in one place. The fans, however, should not be trusted with the details behind the engineering decisions. What did you expect? The sentence you quoted will never be uttered by the engineers - maybe by marketing, but not by engineering.

I don't think it makes sense to only trust the part of a source that supports your statement. Why should the 'fans' have a lot of knowledge concerning the usage of the wing tips to protect the aileron, but be completely wrong about the usage of wing tips to reduce the effect of tip vortices?

@ROIMaison Where do the fans say it is about ground contact protection? No, this document only illustrates nicely that the ailerons double as air brakes (pages 44 and 46). Also, it contains a lot of photos. There is nothing in it which discusses the engineering decisions.

@PeterKämpf I'm sorry, I thought you brought up this source to support your statement about the protecting of ailerons, my bad.

Is this answer suggesting you want to protect the aileron when the wing hits the ground in flight? Isn't that called a "crash" ? When you hit the ground, protecting the ailerons seems to be the least of your problems. The sudden deceleration will cause quite some yaw, and hitting your wingtip means the plane was also rolled (A10 has dihedral) which means the yawing also causes a pitch down. Pitching down at 1 ft above ground is NOT a good thing.

@MSalters: There are different levels of hitting the ground, and the more benign ones would end in a crash if you lose aileron control. A lot of real world experience went into the A-10 (and not the kind of experience you get when watching Hollywood movies). If the wingtip only touches the ground, protecting the ailerons should actually be very high up on your list.

[citation needed]

@MSalters I would think grazing a wingtip would be a very good reason to want aileron authority (i.e. in order to stop grazing said wingtip as quickly as possible.)

@PeterKämpf "The sentence you quoted will never be uttered by the engineers - maybe by marketing, but not by engineering." I wish I had a dollar for every time I've read or heard such a sentence.

I like the answer, and have no reason to doubt it, but I think it would be wise to provide some amount of citation for what you are saying.  If for no other reason than not everyone is familiar with your level of experience...

Hey, we never said the wingtip vortex was the source of all drag. I thought that you can never completely get rid of the vortex, so the best thing to do is to move it somewhere best, such as above the wing so its low pressure (and relatively higher pressure under the wing) cause additional lift. So a drooped wingtip is very strange to me, seemingly doing the exact opposite! But looking at your pictures, they seem to do neither, and instead shift the vortex horizontally away and not so much vertically in ether direction. Which is still strange.

Very very strange answer. Where is the data to confirm such a claim aside from glider mentioning (but notice gliders do have single centered wheel and can roll to the sides once slowed down enough.)

Drooped wing tips can be found on the Cessna 172 - a high wing aircraft not generally associated with extreme low-level maneuvering of a nature that wingtip to ground contact would demand engineering consideration. I can't imagine that the Cessna engineers considered protection of the ailerons in their motivation to fit the 172 with drooped wingtips. However, I can certainly understand such a design choice for its aerodynamic/efficiency benefits.

GVG's Air Vectors (what I would consider a solid, well-researched source) indicates that they are in fact wingtip devices: the wingtips are slightly turned down to improve aileron response and reduce wingtip vortexes that would impair flight efficiency.

I hate to ask it, Peter, but do you have a source from the Athens development where the OEM (Fairchild Republic) designed that wingtip shape for the reasons you stated?

@CarloFelicione No, but the shape can only serve that function. The turned-down part is too small to do anything substantial to the airflow. But it is just big enough to protect the deflected aileron (but not the airbrakes).

@PeterKämpf could you link to theory/tests which show that the turned-down part is too small for substantial beneficial aerodynamic effects?

Partially supporting your answer there's the fact that "the wheels of the main landing gear partially protrude from their nacelles when retracted, making gear-up belly landings easier to control and less damaging". So it was actually designed to land/touch down like a glider

@KennSebesta Size does matter! The effect is proportional to size, so while a small change in the flow will be observed, it will be far too small to show up in performance charts. There are countless reports which will confirm this.

@KennSebesta to complete the comment of Peter: the reduction in induced drag is for sure negligible (a full winglets may reduce drag of some 3% so that small thing might reduce induced drag of some 0.x%) plus there is obviously the relevant increase in the friction drag to be considered. At the end I also suppose that the real job is as explained by Peter plus as a protection for the chaff and flare (which would actually be a job of paramount importance for this kind of airplane). If you want something to read I'd suggest this pdf

@PeterKämpf: what abou this picture?

@sophit Thank you for the link to the McLean article! Figures 4.1 and 4.2 should be required reading for all winglet cult members. The picture shows only that in level flight the turned-down wingtip is not needed. It really is meant for the banked, flying airplane close to the ground.

Nearly certain this is not the answer. Look at glider wing tips design as landing skids (https://nasa.fandom.com/wiki/Wingtip_device?file=Grob_G103C_Glider.jpg) vs that of the A-10 (https://upload.wikimedia.org/wikipedia/commons/e/e2/Thunderbolt_II_080325.jpg). The A-10 couldn't use them as wing-tip skids without cart-wheeling down the runway and the structure is too delicate. I just posted a couple of more possibilities - but I don't have insider knowledge to be certain of why the designers made them this way.

@Jim2B Please read my answer carefully again. It is not about a gear-up landing, and it's not about protecting the split ailerons. It is about maneuvering at very low level while airborne.

@PeterKämpf the only time the wing-tips could possible help with protecting the ailerons would be during a gear-up landing. The problem is that the tips are not long enough to help even under this very rare condition. I added a photo to show the max deflection of the decelerons on the A10 relative to the size of the wing tip devices to my answer.

@Jim2B What your photo shows is the split aileron used for aerobraking. The regular control deflection is much smaller. Please don't contradict things you have no knowledge of.

I thought of another one; if you're installing wingtips and it doesn't matter whether they point up or down, the latter will be less of an obstruction to the view from the cockpit, which is convenient for an aircraft also used for forward observation.
It's an interesting question! I think my answer wasn't mentioned yet?

Aeroalias, ROIMaison, et al all provide excellent explanations and I believe they're correct. However, I think there are two additional possibilities for this wing-tip design.
I have no specific knowledge that indicates the A-10 design considered these issues, but I suspect (educated guess by an Aero & Astro Engineer currently working with the A-10C) one, the other, or both may have been part of the design decision.

Note: the wing tip is very delicate and does not extend far enough to provide support during gear-up landings.
The wing tip devices aren't large enough to protect the decelerons

Or even from more modest aileron deflections

Even if the wing tip devices were larger, the wing dihedral keeps those wing tip devices farther above the ground, even during a gear-up landing, than the ailerons (as shown in the first picture).
Ground Effect
There's another reason for the wingtip "droop" (Hoerner Wing Tip) (note: other references for the Hoerner Wing Tip provide photos or diagrams of several different possible configurations) besides the improved wing efficiency from the winglet effect, namely ground-effect.

For fixed-wing aircraft, ground effect is the reduced aerodynamic drag
that an aircraft's wings generate when they are close to a fixed
surface. During takeoff, ground effect can cause the aircraft to
"float" while below the recommended climb speed. The pilot can then
fly just above the runway while the aircraft accelerates in ground
effect until a safe climb speed is reached.

The A-10 was designed for short takeoff and landing operations. For this type of operation, landing in a short distance is harder than taking off (especially from primitive fields). Ground effect allows a plane to fly slower and remain airborne by providing a region of increased high pressure below the wing when it is close to the ground (low-winged aircraft benefit more from this effect than high-winged aircraft). This allows for lower landing and take-off velocities, which in turn allow for shorter take-off and landing distances.
The Hoerner Wing Tip helps trap this region of high-pressure air under the wing.
Improved Stability
Another possible reason for this configuration, which I discovered while linking references, was apparently the Hoerner Wing Tip helps reduce control instability in dutch roll for wings with dihedral.

This addition was done in order to counteract the dutch roll
characteristic present in the original He 162 design, related to its
wings having a marked dihedral angle.

I wasn't aware of dutch roll instability concerns in the A-10 design before the addition of the Hoerner wing tips but its wing does possess a dihedral. All images that I've seen of the YA-10A wing include the Hoerner wing tip.


text:
Why do trijets (3 rear engines) usually have a T-tail instead of a normal tail?
I'm thinking of trijets like the Tu-154, which has three jet engines in the rear. Not something like L-1011 Tristar which has only one engine at the rear.
Tu-154:
Source: Wikipedia
L-1011 TriStar:
Source: Wikipedia
Why do these kinds of jets have T-tails? Maybe the side jets get in the way of a conventional tail, but it looks to me like you could just raise those engines a little bit and enough room would appear.
My understanding is that T-tails add complexity. The vertical part of the tail has to be strengthened (more weight) since it undergoes more forces. So a conventional tail seems preferable.

related

Some related content-- https://aviation.stackexchange.com/questions/33206/why-does-the-mig-15-have-a-cruciform-tail/67230#67230

A T-tail is not so bad. Its main advantages are:

A smaller vertical tail is required, because the horizontal tail acts like an endplate and enhances the efficiency of the vertical tail.
When the vertical tail is swept, a T-tail will allow to make the horizontal tail smaller as well, because it gains additional lever arm in this configuration.
By designing the junction with the vertical well, the T-tail has less interference drag. It also helps to reduce wave drag, especially when using a well designed Küchemann body (the round, long, spiky thing on the tail junction of a Tu-154) by stretching the structure lengthwise.

As soon as the cruise Mach number demanded tail sweep, the T-tail became the preferred solution. Placing the engines at the rear fuselage put them higher, so FOD became less of a worry, and the wing could be kept clean for maximum lift resulting in shorter runway requirements. In the days before CFD, it was much easier to get the wing right when no pylons with barrels at their ends were sticking from it.
If an engine is mounted near the place where normally the horizontal tail would be, it is much easier to relocate the horizontal to the top of the vertical than trying to join both together. The mass of the engines will require a relatively rearward wing position, so the lever arm of a conventional tail would be rather small. By shifting it up to the top of the swept vertical tail, its lever arm is much larger, making the T-tail especially attractive for configurations with rear-mounted jet engines.
A second reason is better failure tolerance. Note that the T-tail designs are from the early jet age. Their first flight years were: Caravelle 1955, Jet Star 1957 (both not real T-tails, I know, and having little tail sweep), Trident 1962, VC-10 1962, Ilyushin 62 1963, Boeing 727 1963, Tu-134 1963, C-141 1963, Hansa Jet 1964, Tu-154 1968. The designers were afraid an engine failure would also damage the horizontal tail and would turn a survivable accident into a lethal disaster.

I like your comment about engine failure possibly damaging a nearby conventional tail, but the T-154 would seem to shoot itself in the foot by this reasoning, since an uncontained engine failure in its middle engine could disable the entire tail. Also, I can't agree with "its lever arm is doubled". Lever arm is measured from the elevators to the center of mass, or if you're talking about takeoff V1, from the elevators to the wheels. Sweeping the horizontal stabilizer back such that the elevators are a little more rearward does not double the distance to either one.

@DrZ214: 1) The middle engine of the Tu-154 is at the very end of the fuselage. Don't confuse the intake with the engine! 2) It's about vertical tail sweep. Put the horizontal tail at the bottom of the vertical. Now put it on the top. Measure the difference. Can you now agree?

No because again, you are not measuring between the control surface and center of mass, which is somewhere very near the center of the fuselage. While flying, this is the lever arm for elevators. Instead you are picking some other distance, like top of the tail to the bottom. But the pivot point is not the aft of the plane. The pivot point (while flying) is the center of mass, and your lever arm must be measured from there.

@DrZ214: How did I imply that you measure in the side view? I did not. Look at the top view. The center of mass is at the last third of the wing root chord. Draw a realistic horizontal tail at the rear fuselage, and repeat at the tip of the swept vertical tail. Again, measure the difference. It might not be exactly double, but the change in lever arm should be significant.

I did the calculation for Tu-154 and it looks like its T-tail provides a lever arm 1.383 times longer. https://i.sstatic.net/k98CB.png Also, are you sure the Center of Mass is that far back in the Tu-154? I thought CoMs had to be ahead of the Center of Pressure to be stable, and it looks like the last 1/3rd of the wing chord is pretty far back there.

@DrZ214: The wing is swept and produces lift over most of its span. The center of mass looks right. The green tail is really way back - look where Airbus and Boeing place the horizontal. There needs to be some fuselage structure left to attach it to. That was the reason I asked for a realistic horizontal tail. It will be sufficient to look at the horizontal component of the distance - only the part orthogonal to lift counts.

I checked out the Boeing 747 here: https://en.wikipedia.org/wiki/File:B747FAMILYv1.0.png and the A380 here: https://en.wikipedia.org/wiki/File:Airbus_A380-841,_Malaysia_Airlines_AN2208844.jpg Looks to me like it's about right. The back end of my green line doesn't have to be attached to fuselage; it just has to be the outer tip of the swept stabilizer.

@DrZ214: Due to the engines getting in the way a conventional elevator would have to be placed on them and that is quite a bit further forward from the green line you drew.

You want clean, undisturbed air flowing over the control surfaces. You could probably squeeze it in someplace above or below the engines, but it would not be easy without negatively influencing the aerodynamics over the elevators. As you can see from the diagram, you would need to get them pretty far out of the way of the engine cowlings.
If you place either engines or elevators low down along the fuselage they suffer from the disturbed air coming of the wing; this is not desirable. As it stands, the side engines (#1, #3) are already angled for the downwash.
As for the more need for a sturdier rudder, this is partially weighed against that you have a longer moment arm.

(source: state.gov)

Complexity. The worst possible design would be like this: Horizontal and vertical stabilizers are mounted to an engine each...
So you triple the maintenance difficulties you have with a MD-11 or DC-10. And you want an easy access to the engines as possible - which requires enough space.

Would the horizontal stabilizer even work with a jet engine sitting right in front of it? If "no", then the reason for the T-tail isn't "complexity".

@DavidRicherby: Putting it behind the engine is not an option. Putting it on it, similar to how the vertical stabilizer is on the middle engine, is.

@JanHudec: Why isn't that an option?  If anything, the jet exhaust being blown across the stabiliser would seem to increase the effectiveness of the stabiliser and elevators...

@Sean, yes, but 1. the exhaust is hot and 2. you need them to still have enough authority when the engine is out, so it does not allow you to make them smaller anyway.

A horizontal tail that is actually mounted to the fuselage wouldn't work well in a design with jet engines mounted to the fuselage, because it would be directly in the jet exhaust, so it comes down to chosing between a cruciform tail (Caravelle style), or a modified version with the horizontal tail closer to the fuselage, but still above the jet exhaust stream, or a T-tail.   Clearly the T-tail has won out, in the airliner world.  Other answers give some hints as to why.
You might say that if you are going to have the complexity of a horizontal tail mounted to the vertical tail at all, you might as well get the additional benefits of the T-tail.
Yet the T-tail does have some disdvantages over a horizontal tail mounted lower on the vertical tail.
The Falcon family of business jets shows that 3 engines in the rear and a horizontal stabilizer mounted fairly low on the vertical tail can work well- see https://www.dassaultfalcon.com/en/Aircraft/Models/7X/Pages/overview.aspx .
Every aircraft design is influenced by many competing factors and it's hard to set hard-and-fast rules as to what will work best.

you can in theory at least build it like you would with engines mounted in the wing root. Maintenance headache, but no reason the design couldn't work aerodynamically.

Future edit-- add this link to related answer- https://aviation.stackexchange.com/questions/33206/why-does-the-mig-15-have-a-cruciform-tail/67230#67230


text:
Why aren't angle-of-attack data sanity-checked before being fed to other aircraft systems?
For most fixed-wing aircraft, accurate angle-of-attack information is utterly essential for safe flight, due primarily to the risk of stalling the aircraft if the AoA of one or both of its wings becomes excessively high (an additional concern which arises at very high speeds is the possibility of causing an over-g, potentially resulting in structural failure);1 as a result, all airplanes have some form of attack-angle sensor (typically a mechanical vane which aligns itself with the ambient airstream2) which feeds AoA data to the airplane's other systems (usually via the airplane's air-data computers).
Unfortunately, the indispensability of accurate attack-angle data means that, if the AoA data fed to other aircraft systems is not accurate, very bad things happen.  This has caused numerous accidents:

TW843 (L-1011-1, 30 July 1992) - The aircraft's right-hand angle-of-attack sensor, a known temperamental component, failed completely during the beginning of pre-takeoff taxi at JFK, freezing in a position corresponding to an AoA of 26.1° (an angle of attack which is physically impossible for an L-1011 to assume when all of its wheels are on the ground) throughout the aircraft's taxi and takeoff.4,5  This resulted in a false stall warning sounding as soon as the aircraft lifted off,6 leading to a very-high-speed rejected takeoff, excessively-hard touchdown, and runway excursion; structural damage incurred during touchdown resulted in a fire which destroyed the aircraft (although all 292 occupants survived).
QF72 (A330-300, 7 October 2008) - During cruise over the eastern Indian Ocean, one of the aircraft’s three ADIRUs suffered a glitch which corrupted (to a greater or lesser extent) all of the parametric data computed and passed on by same; although this affected measurement and computation of many different aircraft parameters, the important part, for the purposes of this discussion, was that the AoA values transmitted by the malfunctioning ADIRU were contaminated by intermittent (but frequent) spikes to 50.625° (as well as, later on, some spikes to 5.625° or 16.875°), which caused the aircraft’s flight-envelope-protection systems to trigger two large nose-down elevator movements about two-and-two-thirds minutes apart (each time momentarily locking out the pilot’s compensatory nose-up control inputs) in response to a nonexistent stall condition, resulting in numerous severe injuries among the passengers and cabin crew (although all 315 occupants survived).
JT043 / JT610 (737-8, 28-29 October 2018) - During turnaround maintenance at Denpasar on the first day, the aircraft's left-hand AoA sensor was removed due to recurrent malfunctions and replaced with one that had been in storage since being refurbished a year previously.  The replacement sensor proved to have been improperly calibrated during refurbishment, causing it to read 21° higher than the aircraft's actual attack angle; the bad data from the sensor caused the aircraft's stickshaker to activate almost immediately after liftoff on the subsequent flight to Jakarta (it remained on for the remainder of the flight), and, more seriously, interacted with a poorly-designed modification of the aircraft's autotrim system in such a way as to cause the aircraft's horizontal stabiliser to run away in the nose-down direction, resulting in serious control difficulties until the flightcrew were able to stop the trim runaway by disabling the autotrim system.  Maintenance personnel at Jakarta were unable to find a problem with the aircraft on the ground following the flight, and it was released again for flight the following day.  As the miscalibrated AoA vane was not repaired or replaced, the problems from the previous day repeated themselves; the stickshaker activated as the aircraft rotated for takeoff (and once again remained active throughout the flight), and the stabiliser trim began to run away in the nose-down direction as soon as the aircraft's flaps were retracted.7  Unlike what happened the previous day, the flightcrew either forgot or were unable to disconnect the autotrim system before the extreme out-of-trim situation resulted in a loss of control and the aircraft crashed into the Java Sea with the deaths of all 189 on board.

Many of these accidents could have been prevented if some form of sanity checking (a test to detect and screen out nonsensical incoming data) were run on the angle-of-attack data before releasing it to other systems:

An AoA vane that becomes stuck in a fixed position on the ground could be easily detected during takeoff by analysing the data provided by the sensor as soon as the aircraft's airspeed becomes high enough to allow the sensor vanes to operate, and marking as failed any sensor which reports an attack angle that it is not physically possible to assume when the aircraft is on the ground and unrotated.
A sensor reading consistently high or low could be found out by calculating the z-axis load factor the reported angle of attack would create given the aircraft's current IAS, weight, and configuration, and rejecting the data from any sensor that has reported an AoA that is too excessively high or low to produce the observed Gz.
A sensor that is starting to wander away from accuracy would be detected similarly, as its reported AoA would change over time in a manner inconsistent with the changes in the aircraft's airspeed, load factor, weight, and configuration.
A failure that results in the reported attack angle changing unphysically rapidly (for instance, going from +2° to +50° and back to +2° in the space of two seconds8), which could conceivably be produced by certain types of sensor failures,9 but would more likely indicate a problem slightly further downstream, involving the improper processing and conversion of valid sensor data (such as from an ADIRU failure), would be particularly easy to detect, allowing the offending piece of equipment to be rapidly marked as failed and thrown offline.

So why isn't an airplane's raw angle-of-attack data sanity-checked before letting other aircraft systems run wild with it?

1: This does not necessarily require that the aircraft's attack angle be constantly displayed to the pilots; it is often sufficient simply to have the aircraft's systems monitor its AoA, and take action (such as activating the stickshaker) if it becomes excessive.
2: These sensors are typically mounted on the forward fuselage, even though mounting them on the wings - which are, typically, the part of the aircraft that generates most of its lift, and, thus, the part of the aircraft whose angle of attack is actually relevant - would provide more accurate data (because the local airflow - and, thus, local attack angle - over the forward fuselage is generally different from that over the wings; thus, AoA readings from nose-mounted sensors need to be corrected in order to calculate the wings' AoA).3
3: The magnitude and direction of the difference depends, among other things, on the aircraft's airspeed; thus, the process of calculating the wings' attack angle from that of the nose requires valid airspeed data, which can create problems of its own if the aircraft's systems do not actually have access to valid airspeed data.
4: The L-1011 uses an unusual, and ridiculously-complicated, type of AoA sensor, which, rather than a simple mechanical vane, uses magic a hollow tube, pierced with two rows of perforations, which interfaces with a nearby differential-pressure sensor.
5: The data provided by any type of attack-angle sensor will necessarily be meaningless at low airspeeds (such as during taxi and the early portion of takeoff) due to the very low dynamic pressure imposed on the sensors at these low speeds; however, properly-functioning AoA sensors will provide valid data at high airspeeds, even on the ground.
6: Due to the abovementioned unreliability of AoA-sensor data at low speeds, combined with the physical impossibility of stalling an aircraft which is sitting on the ground, the L-1011's stall-warning system is inhibited when the aircraft's air/ground switch is in "ground" mode.
7: The particular subsegment of the 737 MAX (737-7/-8/-9/-10)'s autotrim system which caused the nose-down trim runaways is only active when the flaps are fully retracted.
8: This is not to say that attack angles of this magnitude cannot be experienced in a typical airplane; an airplane's AoA can easily approach 90° in a prolonged stall (especially one that degrades into a flat spin).  However, no airplane larger than a fighter jet - and especially without thrust vectoring - is capable of pitching up by over 45° in the space of a single second; the rotational inertia of a large, enlongated, heavy fuselage is simply much too great for such a rapid, drastic maneouvre to be physically possible.
9: For instance, the L-1011's aforementioned mess of an attack-angle sensor system includes a motor which rotates the external sensor tube to null out the sensor's internal pressure differential, and reads out the local angle of attack by measuring the angle through which the sensor tube is rotated by the internal motor; a short circuit could easily cause the motor to keep running until it hit its mechanical stops in one direction or the other, driving the sensor tube to a very high positive or negative angle, and thereby causing the sensor to report a falsely-high AoA (and possibly one not even in the right direction, at that).

On most systems they are sanity checked against 2 or more other AoA sensors... Are you specifically interested in systems that only have one AoA sensor?

The L-1011's differential pressure sensor is far simpler and more reliable than a vane running dual Rotary Variable Displacement Transducers.  It's somewhat less precise than a vane however so it's not as widely used.  Guys have made AOA sensors for homebuilt aircraft using the same kind of tube with two chambers and rows of holes, hooked up to an airspeed indicator with AOA markings added.

@RonBeyer: QF72 had three ADIRUs, only one of which malfunctioned... didn't help.  JT610 had two AoA vanes, only one of which was miscalibrated, plus load-factor, weight, and configuration data that could also be used to calculate the aircraft's attack angle if the aircraft's designers added a routine to the aircraft's flight control computers to allow them to do so... didn't help.

You could add GXL888T to your list. That was prompted by frozen AoA vanes. That even negated redundancy in the vanes.

@TomMcW: Yes, and I did consider it, but I couldn't see any way that accident could have been prevented or mitigated by sanity-checking the AoA sensor data (the attack angles reported by the sensors were reasonable angles for an airplane to have... they just happened to not be the ones that particular airplane had at that particular moment.

"For most fixed-wing aircraft, accurate angle-of-attack information is utterly essential for safe flight,"   -- faulty premise; this question could be improved by editing to delete this statement.  Thousands or tens of thousands of planes fly every day whose pilots have no direct access to angle-of-attack data.

Ditto what Quiet Flyer said x10.  You are being way to general and inaccurate with your assumptions.  Only modern, complex, (and generally large) aircraft that are heavily reliant on computers have this problem.

And exactly what or how are you going to ‘sanity check’ these systems?  The AoA sensor is just a sensor, converting a physical position into a voltage which another piece of hardware can interpret the results from.  Now some aircraft have multiple AoA vanes so a faulty vane could be detected and isolated if its input data does not concur with other sensors or available information.
Now a lot of this is mitigated through risk assessment ie determining the mean failure rates associated with properly maintained equipment.  I would add the fact that you cite a mere 3 accidents associated with millions of air carrier flight hours over the past 60 years of modern air travel seems to indicate the risk of these accidents to be so small that it may simply be trivial compared with the additional weight gain form additional systems.  Second, this is kind of Monday morning quarterbacking in that it’s easy to say that now AFTER these mishaps occur. Third remember that in the L1011 example these aircraft are products of their time and that aircraft design is an continuously evolving process.  We get better as we go and many time our best lessons were paid for in the blood of the less fortunate.

And yet, despite all the less fortunate blood shed by attack-angle-sensor failures in years past, raw, unchecked AoA data still goes straight to other aircraft systems without first verifying that the attack-angle profile the reported data paints is in fact physically possible.

Yes and no.  Some of the accidents you list are largely to blame on poor maintenance on the part of the operators.  All the system redundancy in the world won’t help a poorly maintained system or one subject to neglect.  But again how exactly do you propose to do a ‘sanity check’ on these systems?

@Sean and how are you going to determine that? And how are you going to know whether the sanity check system is working properly? If the AoA sensor shows the aircraft has an 80% nose up attitude, are you going to decide it's "impossible" because it's outside an "expected range"? When in accidents it might well happen, even if it normally never should.

It is unfortunate that statisticly "mere 3" is considered acceptable.  A "sanity check" is why we have human pilots, competent engineers, and adequate internal and external oversight.  Profit driven people afraid to miss a quarter are not the solution, nor is extensive PC coverage of what has happened.  Aircraft AOA sensor systems simply need to be built better, and hear no evil, see no evil, speak no evil quality control eliminated.  I'd miss 4 quarters (or piss on a spark plug) if it meant a better long term, safer solution.  Unfortunately, a lost reputation cannot be bought back.

But there’s a point where redundancy and safety themselves become redundant, unsafe and expensive.  OEMs strive to walk the fine line between solid safety design practices and excess expense.  And given accident statistics associated with air travel, they do a pretty damn good job!

That they do, could it be better?

@jwenting: An AoA of 80°?  Perfectly possible (although generally undesirable).  An AoA that goes from 5° to 80° in one second and then back to 5° a second after that?  Not so much.  As explained in the latter portion of the question.

@Sean Bit late to the party here, but how would an electrical or mechanical failure mode ever produce that scenario though? Things either seize, fall off, go open circuit or short circuit. I can imagine a dirty pot producing results like you said, but these AoA probes are a sealed unit. The probes are inspected every before flight. Is that not the sanity check?

Most modern systems, such as those found in Air Data System, stall warning, Primary Flight Control Computers, already do bound checks on the measurement values. If an AOA measurement is outside of programmed static thresholds, typically very negative or very large, the system can either inhibit its functionality or vote the measurement out. This is pretty much all a sanity check can do with an isolated measurement (other checks include signal integrity checks, such as parity checks, which are not of our interest here).
For the other checks that you have proposed, they will inevitably include other measurements.
The first problem is, if a measurement disagrees with your model predicted measurement,  how do you tell which is erroneous? Is it the weight, impact pressure, Nz, TAT, or your assumed CL-alpha curve at the measured Mach? If this is a voting scheme, you risk voting off the wrong sensor and exacerbate the problem. The complexity very quickly ramps up as you would need to do cross checks for all the other measurements, and still need to add a large bound to cater for uncertainties.
A better solution is to cross-check against multiple AOA measurements. For example, the A330 uses the median of the three AOA measurements as an integrity source, while the final AOA value used for control law input is the rate-limited average of two offside AOAs.
This comes to the second issue of availability vs integrity, which has been discussed in detail in this answer.

Yes.  Failsafes within failsafes within failsafes soon beget new modes of failure.  For instance, a radio-control quadrotor, powered up indoors for testing, under a metal roof, which notices that it's lost its GPS signal and autocorrects for it by climbing to 100 feet, unsuccessfully.

It is a very good question, in light of what happened with the Lion Air flight 610 crash. Caused by a single mis-calibrated AoA sensor, resulting in 189 deaths.
Single inputs into flight automation are allowed, if the single input can be safely disregarded by the on-board systems. A safety analysis will take into account:

what the consequences of the signal failure are (both fail-to-zero and hard-over fail to a large value);
how safety critical the system is that utilises the signal;
what actions need to be taken (if any), and how fast;
how often the signal fail may occur in the lifetime of the aircraft.

As also discussed in this answer: it must be shown in the safety analysis that the probability of catastrophic failure is less than one in a billion flight hours. This is in an analysis, which can only be statistically verified if the fleet has accumulated several hundred billion flight hours.
And this is where the procedure went astray in the B737MAX certification. The AoA sensor hard-over failure had catastrophic consequences, and was a single input into a flight critical system. Aeronautical Engineering 101: never do this. I am sure that all Boeing engineers involved in the system design were aware of the possible implications, release of the internal communications has shown their concerns and flags.
Yes a sanity check might make sense, particularly in 10/10 hindsight. But in system safety, any action may have an unintended consequence, like in software programming: any statement may introduce a bug, which of course was never foreseen. Even alerting the pilot that something may be wrong with the AoA input may increase pre-TakeOff workload and stress, resulting in more incidents than without the warning. And it may be that the safety analysis makes a compelling case for this side of the coin.
Picture from the crash report
But even a simple sanity check on ground may have pitfalls. Does the AoA sensor always deflect down due to gravity, indicating a negative AoA. or is its position on ground undefined due to small mass unbalance and some allowable friction? The AoA sensors did show a 10° difference on ground, would then all automated systems and stall warnings that use AoA input be disconnected, creating possible other danger scenario's or consequences for dispatch reliability?

The AoA sensor in the B737 max is sanity checked in a sense: in the preceding Lion Air flight SPD and ALT flags appeared on the PFD, and SPEED TRIM and MACH TRIM warning lights illuminated. As a result maintenance performed troubleshooting, also from the crash report:

The AFML entry stated the engineer in Denpasar intended to replace the AOA sensor for trouble shooting due to repetitive problem.

In the balance, the single FCC design of the B737 does seem like a heritage from a bygone era - which it is. AoA sensor failures had not resulted catastrophes in the B737 fleet prior to MAX. From the report page 195:

Erroneous AOA signals are not frequent events. Boeing reported that 25 activations of stick shaker mostly due to AOA failures occurred in 737 aircraft for the past 17 years during more than 240 million flight hours. The MCAS architecture with redundant AOA inputs for MCAS could have been considered but was not required based on the FHA classification of Major.

There is pressure from the airlines to maintain type commonality and limited differences training. All good, but duty of care must be taken to protect us, the passengers, to exclude risks larger than acceptable.

"The AoA sensors did show a 10° difference on ground, would then all automated systems and stall warnings that use AoA input be disconnected, creating possible other danger scenario's or consequences for dispatch reliability?"  1 - I would assume that the thing-to-do for an AoA-disagreement situation would be "don't take off", the same as if any other safety-critical system were inoperative.  2 - the sensors can easily be tested during the takeoff roll, with plenty of time to abort well before V1 if a problem is detected.

The AoA sensor was not identified as being flight critical. Yes indeed the sensors can be tested during TO, they will show a variance from wind direction, altitude etc.

Aircraft usually have "disagree" warnings when redundant sensors do not provide the same reading (within some specified tolerance). This warning may (or may not) inhibit other systems that depend on those sensors.
For instance, the MCAS system in the 737MAX was designed to rely on two AoA sensors. If the two sensors disagreed, the pilots were given a warning and MCAS was inhibited since it wouldn't know which sensor to believe and might cause a crash. Unfortunately, the second AoA input was later changed to an optional feature, and two planes purchased without that feature (despite being physically equipped with multiple AoA sensors) crashed because the failure of the single AoA sensor connected to MCAS was therefore not detected.

Don't think it's quite accurate. The MAX had an option for AOA disagree warning; but it's only a warning. MCAS never depended on dual confirmed AOA; it used to also depend on load factor input.


text:
Does Active Thermal Imaging (NDI method) detect all types of structural faults in modern civil aircraft?
What are the weaknesses of Active thermal imaging in NDI of different composite materials used in aircraft structure ?
How much time and effort does it take to complete a full structure inspection with the latest portable equipment like (C-CheckIR)

No. There is no single NDT method that can identify all types of structural faults in the aircraft.
Active thermography refers to the use of (thermal) exciters like ultrasound, (halogen and xenon) lamps, eddy current transducers and detecting the defects using that fact that the heat produced is different in those areas, usually using infrared cameras.
Though this is a very effective method for detecting defects, especially in composites, there are some disadvantages as well.

The test specimen has to be 'hooked up' to some form of (thermal) loading device. This limits the size of the tested part.
This is an excellent method for detecting subsurface defects; however, as the depth of the defect increases, the defect size has to be larger for detection.
The method requires that the temperature of the region (containing the defect) be uniform. This is a problem especially for materials with high thermal diffusivity like metals, which may require high speed IR cameras.
Though automated algorithms are available for defect detection, most of the (older) methods of thermography depend on the operator for interpretation of results.
The type of excitation technique used can be limited in some cases to prevent damage to the tested part.

So I infer from your statements that the answer is no?  ;-)

@Lnafziger Sadly, yes. After all, there is no magic bullet for solving all problems.

Well, I was just trying to point out that  you don't explicitly answer the question that was asked (especially the title).  You may want to edit slightly just to be clear at the beginning, before going into the why's.


text:
What are the advantages/disadvantages of ultrasonic in aircraft structure inspection?
What i mean by structure is (different composite materials and aluminum).
What are the cons in terms of time required to test the full body of the aircraft (B747 for example) in C and D checks ?

Ultrasonic inspection is perhaps the most widely used inspection procedures in aerospace to detect subsurface anomalies.

Portability- A number of portable ultrasonic equipment are available, which are very simple and are widely used.
The detection depth for flaws is far superior compared to most other NDT methods.
It can be used for detecting flaws even if only one surface of the part is accessible (simply use pulse echo method).
The results are more or less instantaneous- Usually there is no waiting time involved as the results are displayed in a screen.
The time required for preparation is very less; usually the part is cleaned and the couplant applied; after the inspection, it can simply be wiped off.

However, just like any other (NDT) method, this too has its own share of disadvantages:

It is very difficult to use this method for parts which are very thin- for e.g. the aircraft skin because of the dead zone, though this can be overcome by using higher frequencies or immersion technique; in most cases of aircraft checks, this may not be an option. Also, it cannot be used for irregular and non-homogeneous parts (At-least, not the portable methods).
The method does not detect flaws that are oriented parallel to the wave propagation.
The material must be characterized prior to inspection. As the returning  signal is affected by attenuation within the material, comparisons can be made only if a similar (defective and good) specimen of same material is characterized first.
Usually, Ultrasonic method requires a trained operator for interpretation of the results.

As far a full aircraft is considered, only some critical parts are tested using ultrasonic (or other NDT) methods. There is no way to check the entire aircraft except by thorough visual inspection.